Bidirectional control surfaces for use with high speed vehicles, and associated systems and methods

ABSTRACT

Vehicles with bidirectional control surfaces and associated systems and methods are disclosed. In a particular embodiment, a rocket can include a plurality of bidirectional control surfaces positioned toward an aft portion of the rocket. In this embodiment, the bidirectional control surfaces can be operable to control the orientation and/or flight path of the rocket during both ascent, in a nose-first orientation, and descent, in a tail-first orientation for, e.g., a tail-down landing.

CROSS-REFERENCE TO RELATED APPLICATION(S) INCORPORATED BY REFERENCE

The present application is a continuation of U.S. patent applicationSer. No. 14/103,742, filed Dec. 11, 2013, and entitled BIDIRECTIONALCONTROL SURFACES FOR USE WITH HIGH SPEED VEHICLES, AND ASSOCIATEDSYSTEMS AND METHODS, which is a continuation of U.S. patent applicationSer. No. 12/712,083, filed Feb. 24, 2010, and entitled BIDIRECTIONALCONTROL SURFACES FOR USE WITH HIGH SPEED VEHICLES, AND ASSOCIATEDSYSTEMS AND METHODS, which claims priority to U.S. Provisional PatentApplication No. 61/187,268, filed Jun. 15, 2009, and entitled“BIDIRECTIONAL CONTROL SURFACES FOR USE WITH HIGH SPEED VEHICLES, ANDASSOCIATED SYSTEMS AND METHODS,” and which also claims priority to U.S.Provisional Patent Application No. 61/155,115, filed Feb. 24, 2009, andentitled “ROCKETS WITH DEPLOYABLE FLARE SURFACES, AND ASSOCIATED SYSTEMSAND METHODS,” which are incorporated herein by reference in theirentireties.

TECHNICAL FIELD

The present disclosure is directed generally to control surfaces for usewith high speed vehicles, and associated systems and methods.

BACKGROUND

Rocket powered launch vehicles have been used for many years to carryhumans and other payloads into space. Rockets delivered the first humansto the moon, and have launched many satellites into earth orbit,unmanned space probes, and supplies and personnel to the orbitinginternational space station.

Despite the rapid advances in manned and unmanned space flight,delivering astronauts, satellites, and other payloads to space continuesto be an expensive proposition. One reason for this is that mostconventional launch vehicles are only used once, and hence are referredto as “expendable launch vehicles” or “ELVs.” The advantages of reusablelaunch vehicles (RLVs) include the potential of providing low costaccess to space.

Although NASA's space shuttle is largely reusable, reconditioning thereusable components is a costly and time consuming process that requiresextensive ground based infrastructure. Moreover, the additional shuttlesystems required for reentry and landing reduce the payload capabilityof the shuttle. As commercial pressures increase, the need remains forlower-cost access to space. Aspects of the present disclosure aredirected to addressing this challenge.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A is a side elevation view of a representative vehicle havingbidirectional control surfaces configured in accordance with anembodiment of the disclosure, and FIG. 1B is an aft end view of thevehicle of FIG. 1A.

FIGS. 2A-2C are a plan view, inboard end view, and an outboard end view,respectively, of a bidirectional control surface configured inaccordance with an embodiment of the disclosure.

FIGS. 3A and 3B illustrate a representative vehicle during ascent anddescent, respectively, in accordance with an embodiment of thedisclosure.

FIGS. 4A-4D are plan views of bidirectional control surfaces configuredin accordance with other embodiments of the disclosure.

DETAILED DESCRIPTION

The present disclosure is directed generally to bidirectional controlsurfaces for use with rockets and other vehicles that can fly in bothnose-first and tail-first orientations. Several details describingstructures and processes that are well-known and often associated withrockets and aerodynamic control surfaces are not set forth in thefollowing description to avoid unnecessarily obscuring embodiments ofthe disclosure. Moreover, although the following disclosure sets forthseveral embodiments, several other embodiments can have differentconfigurations, arrangements, and/or components than those described inthis section. In particular, other embodiments may have additionalelements, and/or may lack one or more of the elements described belowwith reference to FIGS. 1A-4D.

FIG. 1A is a partially schematic, side elevation view of a vehicle 100having a plurality of bidirectional fins 150 configured in accordancewith an embodiment of the disclosure. FIG. 1B is an aft end view of thevehicle 100 shown in FIG. 1A. Referring to FIGS. 1A and 1B together, thevehicle 100 can be a rocket (e.g., an orbital or suborbital vehicle)that includes a booster or propulsion module 110 carrying a payloadmodule 130. In one embodiment, for example, the vehicle 100 can be areusable launch vehicle that takes advantage of the ability to fly inboth a nose first and tail first direction to recover the vehicle 100 ina vertical, tail first landing. In a particular embodiment, the payloadmodule 130 can be configured to carry cargo and/or crew. In theillustrated embodiment, the payload module 130 has a hemisphericalshape. In other embodiments, however, the payload module 130 can haveother shapes. In still further embodiments, the propulsion module 110can be configured to carry additional rocket stages, such as an upperstage.

The propulsion module 110 can include one or more engines havingcorresponding exhaust nozzles 111 positioned toward an aft portion 101of the vehicle 100. In a particular embodiment, the vehicle 100 includesfive engines, each having a corresponding engine exhaust nozzle 111. Theengines are used during the boost phase to propel the vehicle 100upwardly during ascent. Optionally, some or all of the engine nozzles111 can pivot to provide thrust vectoring to steer the vehicle 100during ascent, either alone or in combination with other control systemsincluding other aerodynamic control systems.

The vehicle 100 can additionally include a deployable aerodynamicsurface or surfaces, such as a deployable flare 140, positioned toward aforward portion 102 of the vehicle 100. The deployable flare 140 can bestowed during ascent and deployed during descent to stabilize and/orslow the vehicle 100 during a tail down descent and landing. In variousembodiments, the vehicle 100 can include deployable flare systems asdescribed in U.S. Provisional Patent Application No. 61/155,115, filedFeb. 24, 2009, and entitled “ROCKETS WITH DEPLOYABLE FLARE SURFACES, ANDASSOCIATED SYSTEMS AND METHODS;” and U.S. Non-provisional PatentApplication Ser. No. 12/712,156, filed Feb. 24, 2010, and entitled“LAUNCH VEHICLES WITH FIXED AND DEPLOYABLE DECELERATION SURFACES, AND/ORSHAPED FUEL TANKS, AND ASSOCIATED SYSTEMS AND METHODS,” both of whichare incorporated herein in their entireties by reference. In theillustrated embodiment, the vehicle 100 can further include a deployablelanding gear 120 (showed stowed in FIGS. 1A and 1B) positioned to allowthe vehicle 100 to land in a tail first or tail down orientation.

Although only illustrative of particular embodiments, the propulsionmodule 110 can have a length L of from about 10 feet to about 50 feet,such as from about 20 feet to about 40 feet, or about 33 feet. Thepropulsion module 110 can also include a cylindrical or circularcross-section having a diameter D of from about five feet to about 20feet, or from about eight feet to about 15 feet, or about 13 feet. Inother embodiments, the vehicle 100 can have other shapes, sizes andoverall dimensions without departing from the present disclosure.

In a particular embodiment, the aft portion 101 of the vehicle 100includes an aft surface 170. In the illustrated embodiment, the aftsurface 170 includes a base region 172 in the proximity of the nozzles111, and a transition region 174. The transition region 174 transitionsbetween the base region 172 and an exterior surface 103 of thepropulsion module 110. In a particular embodiment, the base region 172can be flat, or at least generally flat, and the transition region 174can be curved. For example, in a particular embodiment the transitionregion 174 can have a radius of from about 20 inches to about 50 inches,or about 40 inches. In other embodiments, the base region 172 and/or thetransition region 174 can have other shapes, sizes, and/or dimensions.

In one aspect of this embodiment, the bidirectional fins 150 arepositioned toward the aft portion 101 of the propulsion module 110. Inthe illustrated embodiment, the vehicle 100 includes four fins 150equally spaced around the propulsion module 110, and each of the fins150 is substantially identical, or at least generally similar instructure and function. In other embodiments, however, the vehicle 100can include more or fewer fins positioned at different locations aroundthe propulsion module 110, and one or more of the fins can be differentin structure and/or function.

As described in greater detail below, the bidirectional fins 150 can beused for vehicle guidance and control during both ascent in a nose-firstdirection or orientation, and descent in a tail-first direction. In thisregard, the fins 150 can be operationally coupled to a control system162. The control system 162 can include one or more processors,circuits, and/or mechanisms configured to rotate or pivot the fins backand forth about a pivot axis or hinge line 160 in response to controlsignals received from an on-board guidance system, a remote guidancesystem, and/or computer-readable media. As described in greater detailbelow, the bidirectional fins 150 can pivot together in the samedirection, at the same rate, and/or to the same angle of attack (“α”);or independently (e.g., differentially) with respect to each other indifferent directions, rates, and/or different angles of attack, asrequired to provide the desired vehicle trajectory during ascent and/ordescent. In a particular embodiment, the fins 150 can operate betweenangles of +/−30 degrees. In other embodiments, the fins 150 can pivot toother angles. Further aspects of the fins 150 are described in greaterdetail below.

FIG. 2A is a planform or side elevation view of the fin 150 configuredin accordance with an embodiment of the disclosure. FIG. 2B is aninboard end view of the fin 150, and FIG. 2C is an outboard end view ofthe fin 150. Referring to FIGS. 2A-2C together, the fin 150 includes atip 254 spaced apart from a root 252. In one aspect of this embodiment,the fin 150 has a relatively low aspect ratio (“AR”). For example, thefin 150 can have a span S of from about 15 inches to about 45 inches, orabout 30 inches. The root 252 can have a root chord RC of from about 60inches to about 110 inches, or about 83 inches, and the tip 254 can havea tip chord TC of from about 10 inches to about 30 inches, or about 20inches. As those of ordinary skill in the art will appreciate, theforegoing dimensions are merely representative of certain embodiments ofthe disclosure. The present disclosure is not limited to thesedimensions, and other embodiments can have other dimensions withoutdeparting from the present disclosure.

In another aspect of this embodiment, the fin 150 includes a first orforward edge 256 having relatively little sweep, or no sweep, as definedby a first sweep angle A1 of from about 85 degrees to about 95 degrees,or about 90 degrees. The fin 150 can further include a second or aftedge 258 having a relatively high sweep as defined by a second sweepangle A2 of from about 15 degrees to about 40 degrees, or about 29degrees. In other embodiments, the forward edge 256 and/or the aft edge258 can have other sweep angles. As used herein, in this particularembodiment the term “forward edge” refers to the edge positioned towardthe forward portion 102 of the vehicle, and the term “aft edge” refersto the edge positioned toward the aft portion 101 of the vehicle.

In a particular embodiment, the fin 150 has a symmetrical, or an atleast approximately symmetrical airfoil cross-section. Morespecifically, in the illustrated embodiment the fin 150 has aflat-sided, diamond-shaped cross-section in which the root 252 has amaximum thickness Rt occurring at, or at least proximate to, a midpointMR of the root chord RC. Similarly, the tip 254 has a maximum thicknessTt occurring at, or at least proximate to, the midpoint MT of the tipchord TC. In a particular embodiment, the maximum thickness Rt at theroot chord RC can be from about 6 inches to about 13 inches, or about 9inches, and the maximum thickness Tt at the tip chord TC can be fromabout 1 inch to about 4 inches, or about 2.2 inches. In otherembodiments, the fin 150 can have other symmetric or non-symmetriccross-sections, as well as other maximum chord thicknesses at the rootand/or the tip.

As shown in FIG. 2A, the hinge line 160 is positioned between themidpoint MR of the root 252 and the forward edge 256, and is offset fromthe midpoint a distance HL. In the illustrated embodiment, the distanceHL can be from about 3 inches to about 18 inches, or about 8 inches. Inother embodiments, the hinge line 160 can have other positions relativeto the forward edge 256, the aft edge 258, and/or the midpoint MR of theroot 252.

In a further aspect of this embodiment, the forward edge 256 can have aradius LEr of from about 0.1 inch to about 1 inch, or about 0.25 inch,and the aft edge 258 can have a radius TEr of from about 0.1 inch toabout 1 inch, or about 0.25 inch. In addition, the tip 254 can have aradius Tr of from about 0.1 inch to about 2 inches, or from about 1 inchat the midpoint MT to about 0.25 inch at the forward edge 256 and about0.25 inch at the aft edge 258. Making the tip portion of the fin 150rounded instead of flat can provide gentler stall characteristics. Inother embodiments, however, the forward edge 256, the aft edge 258,and/or the tip 254 can have other shapes, sizes, radiuses and/or otherdimensions. For example, in a particular embodiment the tip 254 can beflat or at least approximately flat.

In particular embodiments, the fin 150 can be manufactured from suitablematerials known in the art, including, for example, suitable metallicmaterials such as aluminum, titanium, and/or steel. In otherembodiments, the fins 150 and/or portions thereof can be manufacturedfrom suitable composite materials, including graphite/epoxy materialsand/or other suitable fiber-reinforced resin materials. Such compositestructures can include, for example, composite sandwich structureshaving a suitable core material covered by a laminated facesheet ofcomposite laminates. In further embodiments, the outer surfaces of allor a portion of the fins 150 can include suitable layers and/or coatings(e.g., ablative coatings) for dealing with the potentially hightemperatures experienced during ascent and/or descent of the vehicle 100(FIG. 1A).

As discussed above, the fin 150 can be implemented to provide guidanceand control on a vehicle (e.g., a rocket) that flies in a firstdirection (e.g., nose first or forward) during ascent and a seconddirection (e.g., tail first or aft-first) during descent. One feature ofthe fin 150 is that when the vehicle is flying in an ascent direction,as indicated by arrow A, the fin 150 provides a relatively high changein lift force as the angle of attack (α) of the fin 150 changes. Putanother way, the fin 150 demonstrates a relatively high lift slopeduring ascent, with lift stall occurring at an angle of attack α of fromabout 8 degrees to about 13 degrees, or at about 10 degrees or more. Asused herein, the term “lift slope” refers to the slope of a curvedescribing the lift, or more specifically the coefficient of lift C_(L),of the fin 150 as a function of angle of attack, α. When the vehicle isflying in a descent direction, however, as indicated by arrow D, the fin150 demonstrates a relatively low lift slope with a peak liftcoefficient C_(L) of at least about 1. Moreover, during descent the fin150 of this embodiment stalls at angles of attack α greater than about12 degrees to about 18 degrees, or greater than about 15 degrees.Accordingly, for reasons discussed in more detail below, in theillustrated embodiment the fins 150 are configured to provide arelatively aggressive lift curve during ascent in a nose firstdirection, and a relatively gradual lift curve, with a relatively highlift peak, during descent in a tail first direction.

In another aspect of the illustrated embodiment, the fin 150 maintains acenter of pressure location during all phases of flight that isrelatively close to the actuator hinge line 160. This minimizes or atleast reduces the torques required to pivot the fin 150 relative to itsneutral state and achieve the desired angles of attack. A further aspectof the fin 150 is that it is configured to operate in a flight regime orenvelope including both subsonic and supersonic flight, includingsupersonic flight at a mach number of about four.

As mentioned above, in a particular embodiment the fin 150 can have asymmetrical, or an at least approximately symmetrical airfoil shape(e.g., a diamond-shape or a “double wedge” supersonic airfoil shape). Asymmetric airfoil can facilitate predictable behavior duringbidirectional flight, and results in the maximum thickness Rt of theroot 252 being positioned relatively close to the pivot axis or hingeline 160.

During ascent in the direction of arrow A, the forward edge 256 is the“leading edge” and the planform of the fin 150 represents a relativelylow aspect ratio AR lifting surface having a non-swept (or very lowsweep) leading edge. In this particular embodiment, this planformcreates a moderate to high lift curve slope with stall occurring beyonda desired angle of attack, such as about 10 degrees. During descent inthe direction of the arrow D, the aft edge 258 becomes the “leadingedge,” and the planform represents a relatively low aspect ratio ARlifting surface having a leading edge that is highly swept at an angleof, e.g., about 60 degrees relative to the airflow. During descent, thishighly swept, low aspect ratio AR planform can provide a relatively lowlift curve slope with maximum lift occurring at relatively high anglesof attack across the entire flight regime. Moreover, during descent thisfin planform can provide a lift stall that occurs at angles of attack ofabout 20 degrees at subsonic speeds, and at more than about 45 degreesat supersonic speeds. During descent, the maximum coefficient of liftcan be at least about 1.0 (for subsonic flight) with peak coefficient oflift values closer to about 1.5 during supersonic flight.

A further aspect of the illustrated fin planform is that during bothascent and descent, the center of pressure location is relatively wellbounded throughout the range of angles of attack. This can minimize orat least reduce the torque required to control the fin 150. Moreover,with this fin planform many of the aerodynamic conditions that result inrelatively high stresses occur when the center of pressure is very closeto the hinge line 160. Although the center of pressure position can, insome embodiments, vary to a greater degree, this is expected to occurduring fin maneuvers and/or aerodynamic conditions that result inrelatively low stresses.

FIG. 3A is a partially schematic, side elevation view of an embodimentof the vehicle 100 during its ascent, as indicated by arrow A. Duringthe ascent or boost phase, the deployable flare 140 is stowed and isaccordingly positioned flat against and/or flush with the externalsurface 103 of the vehicle 100. Moreover, during the ascent phase thelanding gear 120 (FIGS. 1A and 1B) can be stowed.

During boost phase, the fins 150 provide a stabilizing effect as theytend to move the center of pressure aft of the vehicle center ofgravity. In certain embodiments, the degree of stabilization provided bythe fins 150 can be directly proportional to the curve of the lift slopeof the fins and, accordingly, the higher the lift slope the greater thedegree of stabilization. In certain embodiments, the magnitude of thelift generated by the fins 150 may not be as important as the slope ofthe lift curve or the need for the lift curve to remain linear, or atleast approximately linear, over the operational angle of attack range.As mentioned above, the fins 150 can also pivot to help actively guideand control the vehicle during ascent.

FIG. 3B illustrates the vehicle 100 during its descent phase, asindicated by arrow D. During descent, the deployable flare 140 can bedeployed by, for example, pivoting the flare 140 so that it expandsoutwardly from the external surface 103. As discussed above, thisconfiguration is expected to slow and help stabilize the vehicle 100during descent. For example, by deploying the flare 140 the center ofpressure acting on the vehicle 100 can shift upwardly (e.g., above thevehicle center of gravity) so that gravitational forces acting on thevehicle 100 tend to stabilize perturbations that may be caused byaerodynamic forces acting on the vehicle 100.

During descent of the vehicle 100, the engines are off and no longerthrusting in most, if not all embodiments. In certain embodiments, theengines will remain off and non-thrusting until just prior to touch downof the vehicle 100 in a tail-first orientation at the landing site. As aresult, the fins 150 are the dominant aerodynamic control surfaces andthe only means, or at least the predominant means, for steering thevehicle 100 during descent.

During descent, the fins 150 are positioned towards the direction offlight and can thus destabilize the vehicle. In certain embodiments,however, having a relatively gentle lift curve can minimize, or at leastreduce, the aerodynamic destabilization effect of the fins 150 duringdescent. However, because the fins 150 are used for vehicle guidance andcontrol during descent, it is also desirable for the fins 150 to be ableto provide sufficiently high levels of peak lift. This peak lift willenable the fins 150 to orient the vehicle to relatively large angles ofattack when needed during descent.

In another aspect of the illustrated embodiment, the fins 150 arelocated relatively far aft on the vehicle 100. This can maximize, or atleast increase, the ability of the fins 150 to stabilize the vehicle 100during ascent and control the vehicle 100 during descent. As discussedabove with reference to FIGS. 1A and 1B, the aft surface 170 of thevehicle 100 can be rounded in the transition region 174 between therelatively flat base region 172 and the external surface 103 of thepropulsion module 110. As a result, moving the fins 150 aft produces aslight overhang gap 390 between the inboard tip of the aft edge 258(FIG. 2A) and the transition region 174 of the aft surface 170. It isexpected, however, that the overhang gap 390 will not negatively affectoperation of the fins 150 over the flight regime and mission, includingboth forward travel during ascent and aft travel during descent.

During descent, the fins 150 are positioned sufficiently behind a bowshock 380. The relatively flat base region 172 of the aft surface 170tends to move the bow shock 380 outwardly in front of the aft surface170 during descent of the vehicle 100. As a result, the fins 150 arepositioned generally aft or behind the bow shock 380, which can avoid orat least reduce shocks and other high loads on the fins 150 duringdescent.

There are various aspects of the fin design that are expected to providefavorable characteristics for use with a reusable launch vehicle thatcan ascend in a nose-first direction and descend in a tail-firstdirection. For example, the fins 150 are relatively small and, as aresult, remain positioned behind the bow shock 380 during both descentand ascent. As discussed above, this can prevent or at least reduce thelikelihood that shocks will directly impinge on the fin surface andcreate high local loads or unsteady, buffeting loads during flight. Therelatively short fin span S (FIG. 2A) also facilitates working aroundthe vehicle and performing ground maneuvers such as vehicle lifting,rotation, and/or transportation with conventional on-site equipment.

FIGS. 4A-4D are a series of side elevation views of portions of launchvehicles 400 a-d having bidirectional control surfaces or fins 450 a-dconfigured in accordance with other embodiments of the disclosure.Referring first to FIG. 4A, the fin 450 a is at least generally similarin structure and function to the fin 150 described in detail above.However, in the illustrated embodiment the fin 450 a includes an aftedge 458 having a non-swept inboard portion 458 a-1 and a highly sweptoutboard portion 458 a-2 (e.g., an outer one-half portion). In oneaspect of this embodiment, having the aft edge 458 with a straightinboard portion 458 a-1 and a highly swept outboard portion 458 a-2 mayresult in a fin with earlier stall characteristics than the fin 150described in detail above.

FIG. 4B illustrates a fin 450 b having a relatively low or moderatelyswept forward edge 456 b and a relatively highly swept aft edge 458 b.Referring next to FIG. 4C, the fin 450 c has a symmetrical, or an atleast approximately symmetrical planform in which both a forward edge456 c and an aft edge 458 c are moderately to highly swept. Referringnext to FIG. 4D, in this embodiment the fin 450 d has a highly sweptforward edge 456 d and a non-swept or relatively low sweep aft edge 458d. In this particular embodiment, however, the fin 450 d can rotate afull 360 degrees about a hinge line 460 d to that the planform can beoptimized for the direction of flight. For example, in a particularembodiment the fin 450 d can be oriented as shown by the solid line inFIG. 4D for ascent, and then rotated 180 degrees about the hinge line460 d to the position shown by the dotted line in FIG. 4D for descent.Although the fins illustrated in FIGS. 4A and 4D can have symmetricalcross-sections (e.g., diamond-shaped cross-sections), in otherembodiments these fin configurations and variations thereof can havenon-symmetrical cross-sections.

From the foregoing, it will be appreciated that specific embodiments ofthe disclosure have been described herein for purposes of illustration,but that the disclosure may include other embodiments as well. Forexample, the bidirectional control surfaces 150 can have other shapesand/or arrangements that are different than those shown and describedabove depending on the type of rocket, mission, etc. Certain aspects ofthe disclosure described in the context of particular embodiments may becombined or eliminated in other embodiments. Further, while advantagesassociated with certain embodiments have been described in the contextof those embodiments, other embodiments may also exhibit such advantagesand not all embodiments need necessarily exhibit such advantages to fallwithin the scope of the invention. Accordingly, the invention is notlimited, except as by the appended claims.

1-22. (canceled)
 23. A propulsion module for a space launch vehicle, thepropulsion module comprising: a forward portion spaced apart from an aftportion along a longitudinal axis of the propulsion module; a rocketengine comprising at least one rocket exhaust nozzle, the rocket exhaustnozzle positioned toward the aft portion and configured to providethrust to launch the propulsion module in an ascent orientation in whichthe forward portion leads the aft portion; and a plurality ofaerodynamic control surfaces, wherein each of the aerodynamic controlsurfaces is configured to pivot about a corresponding pivot axisextending outwardly from the longitudinal axis to control the propulsionmodule when the propulsion module is flying in a descent orientation inwhich the aft portion leads the forward portion.
 24. The propulsionmodule of claim 23, further comprising a cylindrical portion positionedbetween the forward portion and the aft portion, wherein the pluralityof aerodynamic control surfaces is spaced around the cylindricalportion.
 25. The propulsion module of claim 23 wherein at least one ofthe aerodynamic control surfaces has a symmetrical cross-section. 26.The propulsion module of claim 23, further comprising a landing gearpositioned toward the aft portion, wherein the landing gear isconfigured to be stowed during at least a portion of a flight regime,and wherein the landing gear is further configured to be deployed tosupport the propulsion module in a vertical orientation.
 27. Thepropulsion module of claim 23 wherein the rocket engine is configured tobe restarted when the propulsion module is flying in the descentorientation.
 28. The propulsion module of claim 23 wherein the pluralityof aerodynamic control surfaces is positioned toward the aft portion.29. The propulsion module of claim 23 wherein the plurality ofaerodynamic control surfaces is positioned toward the forward portion.30. The propulsion module of claim 23 wherein the rocket exhaust nozzleis configured to pivot to provide thrust vectoring when the propulsionmodule is flying in the descent orientation.
 31. The propulsion moduleof claim 23 wherein the forward portion is configured to carry a payloadmodule.
 32. A rocket comprising: a first end portion; a second endportion positioned opposite the first end portion; at least one rocketexhaust nozzle positioned toward the second end portion and configuredto provide thrust when the rocket is flying in an ascent orientation inwhich the first end portion leads the second end portion, and when therocket is flying in a descent orientation in which the second endportion leads the first end portion; a plurality of aerodynamic controlsurfaces, wherein each aerodynamic control surface includes a rootportion positioned toward an exterior surface of the rocket and a tipportion positioned outwardly from the root portion, wherein eachaerodynamic control surface is configured to pivot about a correspondingpivot axis extending between the root portion and the tip portion; and acontrol system that pivots at least one of the aerodynamic controlsurfaces about its corresponding pivot axis to change the angle ofattack of the at least one aerodynamic control surface and control therocket in the descent orientation.
 33. The rocket of claim 32 whereinthe control system is further configured to pivot at least one of theaerodynamic control surfaces to steer the rocket in the ascentorientation.
 34. The rocket of claim 32 wherein each of the aerodynamiccontrol surfaces is configured to pivot between +/−30 degrees about itscorresponding pivot axis.
 35. The rocket of claim 32, further comprisinga stowable landing gear positioned toward the second end portion,wherein the landing gear is configured to be deployed to support therocket upon landing in a vertical orientation.
 36. A launch vehiclesystem comprising: a rocket stage having a forward end portion and anaft end portion positioned opposite the forward end portion; a rocketengine coupled to a rocket exhaust nozzle, the rocket exhaust nozzlepositioned toward the aft end portion; at least one moveable controlsurface carried by the rocket stage, wherein the control surfaceincludes a root portion and a tip portion positioned outwardly from theroot portion; and a controller configured to execute instructions that,when executed, cause the launch vehicle system to perform a methodcomprising: operating the rocket engine to provide thrust for launchingthe rocket stage in an ascent phase in which the forward end portionleads the aft end portion; terminating the ascent phase; pivoting thecontrol surface about an axis extending outwardly from the root portionof the control surface toward the tip portion of the control surface tocontrol the rocket stage during a descent phase in which the second endportion leads the first end portion; and operating the rocket engine toprovide thrust for landing the rocket stage in a vertical orientation ata landing site.
 37. The launch vehicle system of claim 36 wherein themethod further comprises deploying a stowable landing gear to supportthe rocket stage in the vertical orientation at the landing site. 38.The launch vehicle system of claim 36 wherein the method furthercomprises: turning off the rocket engine after the ascent phase; andreigniting the rocket engine prior to landing the rocket stage in thevertical orientation at the landing site.
 39. The launch vehicle systemof claim 36 wherein the method further comprises aerodynamicallycontrolling the rocket stage in the ascent phase by moving the controlsurface about the axis.
 40. The launch vehicle system of claim 36wherein the method further comprises reusing at least a portion of therocket stage in a subsequent launch mission.